The present invention relates to gas turbine blades for a turbomachine. More particularly, the invention relates to cooling circuits for such blades.
It is known that the moving blades of a turbomachine gas turbine, and in particular of the high pressure turbine, are subjected to very high temperatures from the combustion gases when the engine is in operation. These temperatures reach values that are well above those that can be withstood without damage by the various parts that come into contact with said gases, thereby limiting the lifetime of said parts.
It is also known that raising the temperature of the gas in the high pressure turbine increases turbomachine efficiency, i.e. the ratio of thrust from the engine over the weight of an airplane propelled by said turbomachine. Consequently, efforts are made to provide turbine blades that are capable of withstanding ever-higher temperatures.
In order to solve this problem, it is general practice to provide such blades with cooling circuits seeking to reduce their temperature. By means of such circuits, cooling air which is generally inserted into the blade via its root travels along the blade following a path formed by cavities made in the blade, and is then ejected via orifices that open out into the surface of the blade.
Thus, French patent No. 2 765 265 proposes a set of turbine blades each cooled by a helical strip, by means of an impact system, and by means of a system of bridges. Although the cooling appears to be satisfactory, such circuits are complex to make and it is found that the heat exchange produced by the flow of cooling air is not uniform, thereby leading to temperature gradients that penalize the lifetime of the blade.